Gas turbine engine with a double wall core casing

ABSTRACT

A gas turbine engine includes an engine core including: a compressor system including first, lower pressure, compressor, and a second, higher pressure, compressor; and an outer core casing surrounding the compressor system and including a first flange connection arranged to allow separation of the outer core casing at an axial position of the first flange connection, wherein the first flange connection is the first flange connection that is downstream of an axial position defined by the axial midpoint between the mid-span axial location on the trailing edge of the most downstream aerofoil of the first compressor and the mid-span axial location on the leading edge of the most upstream aerofoil of the second compressor; a nacelle surrounding the engine core and defining a bypass duct between the engine core and the nacelle; wherein an axial midpoint of the radially outer edge is defined as the fan OGV tip centrepoint.

The present disclosure relates to a gas turbine engine for an aircraft,and more specifically to a gas turbine engine with specified relativecomponent positions.

The skilled person would appreciate that simply scaling up an engine mayintroduce problems such as increased stress, strain and/or bendingmoment on parts of the engine, and/or on a wing of the aircraft to whichit is mounted. Reconsideration of engine parameters may therefore beappropriate.

For example, the skilled person would appreciate that, if the fan sizeof a gas turbine engine is increased, bending loads on the engine coremay be deleteriously increased. A redesign of the engine core and/or ofsupporting components of the gas turbine engine may therefore beappropriate.

According to a first aspect, there is provided a gas turbine engine foran aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising a first flange connection arranged to allow            separation of the outer core casing at an axial position of            the first flange connection, the first flange connection            having a first flange radius, wherein the first flange            connection is the first flange connection that is downstream            of an axial position defined by the axial midpoint between            the mid-span axial location on the trailing edge of the most            downstream aerofoil of the first compressor and the mid-span            axial location on the leading edge of the most upstream            aerofoil of the second compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially outer edge is defined as the fan OGV tip        centrepoint.

A fan OGV tip position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

is equal to or less than 1.8.

The fan OGV tip position ratio may be greater than or equal to 0.6.

The fan OGV root position ratio may be less than or equal to 1.20, andoptionally less than or equal to 1.00.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally greater than 300 cm and less than or equal to 380cm.

The fan diameter may be between 330 cm and 380 cm, and optionallybetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The engine may further comprise a gearbox that receives an input from acore shaft and outputs drive to the fan so as to drive the fan at alower rotational speed than the core shaft. A gear ratio of the gearboxmay be between 3.1 and 4.0.

The first flange connection may be at, or axially downstream of, aleading edge of the most upstream aerofoil of the second compressor.

The first flange connection may be at, or axially upstream of, a leadingedge of the most upstream aerofoil of the second compressor.

The engine may comprise:

-   -   a first turbine, and a first core shaft connecting the first        turbine to the first compressor; and    -   a second turbine and a second core shaft connecting the second        turbine to the second compressor.

The second turbine, second compressor, and second core shaft may bearranged to rotate at a higher rotational speed than the first coreshaft.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

The fan OGV tip position to fan diameter ratio may be greater than orequal to 0.095.

An axial midpoint of the radially inner edge may be defined as the fanOGV root centrepoint. A fan OGV root position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 2.6.

The fan OGV root position ratio may be greater than or equal to 0.8.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

The gas turbine engine may further comprise a front mount arranged to beconnected to a pylon. A front mount position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}} \\{{flange}\mspace{14mu}{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.18.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{20mu}{diameter}}$

may be less than or equal to 0.145.

The engine core may further comprise an inner core casing providedradially inwardly of the compressor blades of the compressor system. Theinner core casing and the outer core casing may define a core workinggas flow path therebetween. A gas path radius may be defined as theouter radius of the core gas flow path at the axial position of thefirst flange connection, and a gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

may be equal to or greater than 1.10 and less than or equal to 2.0.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

may be equal to or greater than 0.125 and less than or equal to 0.17.

A fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

may be equal to or less than 19.0 mm/lb.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising a first flange connection arranged to allow            separation of the outer core casing at an axial position of            the first flange connection, the first flange connection            having a first flange radius, wherein the first flange            connection is the first flange connection that is downstream            of an axial position defined by the axial midpoint between            the mid-span axial location on the trailing edge of the most            downstream aerofoil of the first compressor and the mid-span            axial location on the leading edge of the most upstream            aerofoil of the second compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially inner edge is defined as the fan OGV root        centrepoint.

A fan OGV root position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{11mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

is equal to or less than 2.6.

The fan OGV root position ratio may be greater than or equal to 0.8.

The fan OGV root position ratio may be less than or equal to 2.00, andoptionally less than or equal to 1.10.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally greater than 300 cm and less than or equal to 380cm.

The fan diameter may be between 330 cm and 380 cm, and optionallybetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The gas turbine engine may further comprise a gearbox that receives aninput from a core shaft and outputs drive to the fan so as to drive thefan at a lower rotational speed than the core shaft. A gear ratio of thegearbox may be between 3.1 and 4.0.

The first flange connection may be at, or axially downstream of, aleading edge of the most upstream aerofoil of the second compressor.

The first flange connection may be at, or axially upstream of, a leadingedge of the most upstream aerofoil of the second compressor.

The gas turbine engine may comprise:

-   -   a first turbine, and a first core shaft connecting the first        turbine to the first compressor; and    -   a second turbine and a second core shaft connecting the second        turbine to the second compressor.

The second turbine, second compressor, and second core shaft may bearranged to rotate at a higher rotational speed than the first coreshaft.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

The fan OGV root position to fan diameter ratio may be greater than orequal to 0.12.

An axial midpoint of the radially outer edge may be defined as the fanOGV tip centrepoint. A fan OGV tip position ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.8.

The fan OGV tip position ratio may be greater than or equal to 0.6.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

The gas turbine engine may further comprise a front mount arranged to beconnected to a pylon. A front mount position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.18.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.145.

The engine core may further comprise an inner core casing providedradially inwardly of the compressor blades of the compressor system. Theinner core casing and the outer core casing may define a core workinggas flow path therebetween. A gas path radius may be defined as theouter radius of the core gas flow path at the axial position of thefirst flange connection. A gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

may be equal to or greater than 1.10 and less than or equal to 2.0.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

may be equal to or greater than 0.125 and less than or equal to 0.17.

A fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

may be equal to or less than 19.0 mm/lb.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising an engine core comprising a compressor systemwith compressor blades comprising respective aerofoils, the compressorsystem comprising a first, lower pressure, compressor, and a second,higher pressure, compressor, an outer core casing surrounding thecompressor system. The gas turbine further comprises a fan locatedupstream of the engine core, the fan comprising a plurality of fanblades and having a fan diameter.

The outer core casing comprises a first flange connection arranged toallow separation of the outer core casing at an axial position of thefirst flange connection, the first flange connection having a firstflange radius, wherein the first flange connection is the first flangeconnection that is downstream of an axial position defined by the axialmidpoint between the mid-span axial location on the trailing edge of themost downstream aerofoil of the first compressor and the mid-span axiallocation on the leading edge of the most upstream aerofoil of the secondcompressor, and a front mount arranged to be connected to a pylon.

A front mount position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{connection}} \\{{and}\mspace{14mu}{the}\mspace{11mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

is equal to or less than 1.18.

The front mount position ratio may be greater than or equal to 0.65.

The front mount position ratio may be less than or equal to 1.10, andoptionally less than or equal to 1.00.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally greater than 300 cm and less than or equal to 380cm.

The fan diameter may be between 330 cm and 380 cm, and optionallybetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The engine may further comprise a gearbox that receives an input from acore shaft and outputs drive to the fan so as to drive the fan at alower rotational speed than the core shaft. A gear ratio of the gearboxmay be between 3.1 and 4.0.

The first flange connection may be at, or may be axially downstream of,a leading edge of the most upstream aerofoil of the second compressor.Alternatively, the first flange connection may be axially upstream of aleading edge of the most upstream aerofoil of the second compressor.

The engine may comprise:

-   -   a first turbine, and a first core shaft connecting the first        turbine to the first compressor; and    -   a second turbine and a second core shaft connecting the second        turbine to the second compressor.

The second turbine, second compressor, and second core shaft may bearranged to rotate at a higher rotational speed than the first coreshaft.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.145.

The front mount position to fan diameter ratio may be greater than orequal to 0.07.

The front mount may be a core mount.

The gas turbine engine may further comprise:

-   -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially inner edge is defined as the fan OGV root        centrepoint.

A fan OGV root position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 2.6.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

The gas turbine engine may further comprise:

-   -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially outer edge is defined as the fan OGV tip        centrepoint.

A fan OGV tip position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.8.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

The engine core may further comprise an inner core casing providedradially inwardly of the compressor blades of the compressor system. Theinner core casing and the outer core casing may define a core workinggas flow path (A) therebetween. A gas path radius may be defined as theouter radius of the core gas flow path (A) at the axial position of thefirst flange connection. A gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

may be equal to or greater than 1.10 and less than or equal to 2.0.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

may be equal to or greater than 0.125 and less than or equal to 0.17.

A fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

may be equal to or less than 19.0 mm/lb.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising:        -   a first flange connection arranged to allow separation of            the outer core casing at an axial position of the first            flange connection, the first flange connection having a            first flange radius, wherein the first flange connection is            the first flange connection that is downstream of an axial            position defined by the axial midpoint between the mid-span            axial location on the trailing edge of the most downstream            aerofoil of the first compressor and the mid-span axial            location on the leading edge of the most upstream aerofoil            of the second compressor; and    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades;    -   wherein a fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

is equal to or less than 19.0 mm/lb.

The fan blade mass ratio may be equal to or greater than 5 mm/lb.

A blade set mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{total}\mspace{14mu}{mass}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{plurality}\mspace{14mu}{of}\mspace{14mu}{fan}\mspace{14mu}{blades}}$

may be in the range between 0.95 mm/lb and 0.35 mm/lb.

Each of the fan blades may be at least partly formed from a metallicmaterial. The metallic material may be titanium or aluminium lithiumalloy.

Each of the fan blades may be formed at least partly from a compositematerial.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally may be greater than or equal to 300 cm and lessthan or equal to 380 cm.

The fan diameter may be between 330 cm and 380 cm, and optionally may bebetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The mass of each fan blade may be in a range between 20 lb and 70 lb.

The gas turbine engine may further comprise a gearbox that receives aninput from a core shaft and outputs drive to the fan so as to drive thefan at a lower rotational speed than the core shaft.

A gear ratio of the gearbox may be between 3.1 and 4.0.

The first flange may be at, or axially downstream of, a leading edge ofthe most upstream aerofoil of the second compressor. Alternatively, thefirst flange may be at, or axially upstream of, a leading edge of themost upstream aerofoil of the second compressor.

The gas turbine engine may further comprise: a first turbine, and afirst core shaft connecting the first turbine to the first compressor;and a second turbine and a second core shaft connecting the secondturbine to the second compressor. The second turbine, second compressor,and second core shaft may be arranged to rotate at a higher rotationalspeed than the first core shaft.

The engine core may further comprise an inner core casing providedradially inwardly of the compressor blades of the compressor system, theinner core casing and the outer core casing defining a core working gasflow path therebetween. A gas path radius may be defined as the outerradius of the core gas flow path at the axial position of the firstflange, and a gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

may be equal to or greater than 1.10 and less than or equal to 2.0.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

may be equal to or greater than 0.125 and less than or equal to 0.17.

The gas turbine engine may further comprise a front mount arranged to beconnected to a pylon.

A front mount position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.18.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.145.

The front mount may be a core mount.

The gas turbine engine may further comprise: a nacelle surrounding theengine core and defining a bypass duct between the engine core and thenacelle; and a fan outlet guide vane (OGV) extending radially across thebypass duct between an outer surface of the engine core and the innersurface of the nacelle, the fan OGV having a radially inner edge and aradially outer edge, wherein an axial midpoint of the radially inneredge is defined as the fan OGV root centrepoint.

A fan OGV root position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 2.6.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

A fan OGV tip position ratio of:

$\frac{\begin{matrix}{\mspace{20mu}{{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.8.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

According to a further aspect, there is provided gas turbine engine foran aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor;        -   an inner core casing provided radially inwardly of the            compressor blades of the compressor system; and        -   an outer core casing surrounding the compressor system, the            inner core casing and the outer core casing defining a core            working gas flow path therebetween, the outer core casing            comprising:        -   a first flange connection arranged to allow separation of            the outer core casing at an axial position of the first            flange connection, the first flange connection having a            first flange radius, wherein the first flange connection is            the first flange connection that is downstream of an axial            position defined by the axial midpoint between the mid-span            axial location on the trailing edge of the most downstream            aerofoil of the first compressor and the mid-span axial            location on the leading edge of the most upstream aerofoil            of the second compressor; and    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   wherein a gas path radius is defined as the outer radius of the        core gas flow path at the axial position of the first flange        connection, and a gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

is equal to or greater than 1.10.

The gas path ratio may be equal to or greater than 1.50.

The gas path ratio may be less than or equal 2.0.

The outer core casing may comprise a first outer core casing and asecond outer core casing, the first outer core casing being providedradially inwardly of the second outer core casing. The first flangeconnection may be provided on the second outer core casing. The gas pathradius may be defined as the radius of a radially inner surface of thefirst outer core casing.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally may be greater than 300 cm and less than or equalto 380 cm.

The fan diameter may be between 330 cm and 380 cm, and optionally may bebetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The engine may further comprise a gearbox that receives an input from acore shaft and outputs drive to the fan so as to drive the fan at alower rotational speed than the core shaft.

A gear ratio of the gearbox may be between 3.1 and 4.0.

The first flange connection may be at, or axially downstream of, aleading edge of the most upstream aerofoil of the second compressor.Alternatively, the first flange connection may be at, or axiallyupstream of, a leading edge of the most upstream aerofoil of the secondcompressor.

The engine may comprise: a first turbine and a first core shaftconnecting the first turbine to the first compressor; and a secondturbine and a second core shaft connecting the second turbine to thesecond compressor. The second turbine, second compressor, and secondcore shaft may be arranged to rotate at a higher rotational speed thanthe first core shaft.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

may be equal to or greater than 0.125 and less than or equal to 0.17.

A fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

may be equal to or less than 19.0 mm/lb.

A blade set mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{total}\mspace{14mu}{mass}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{plurality}\mspace{11mu}{of}\mspace{14mu}{fan}\mspace{11mu}{blades}}$

may be in the range between 0.95 mm/lb and 0.35 mm/lb.

The gas turbine engine may further comprise a front mount arranged to beconnected to a pylon.

A front mount position ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{20mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.18.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{20mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{20mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.145.

The front mount may be a core mount.

The gas turbine engine may further comprise: a nacelle surrounding theengine core and defining a bypass duct between the engine core and thenacelle; and a fan outlet guide vane (OGV) extending radially across thebypass duct between an outer surface of the engine core and the innersurface of the nacelle, the fan OGV having a radially inner edge and aradially outer edge. An axial midpoint of the radially inner edge may bedefined as the fan OGV root centrepoint.

A fan OGV root position ratio of:

$\frac{\begin{matrix}{\mspace{20mu}{{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 2.6.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{20mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{20mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

A fan OGV tip position ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{20mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.8.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{20mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{20mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising:        -   a first flange connection arranged to allow separation of            the outer core casing at an axial position of the first            flange connection, the first flange connection having a            first flange radius, wherein the first flange connection is            the first flange connection that is downstream of an axial            position defined by the axial midpoint between the mid-span            axial location on the trailing edge of the most downstream            aerofoil of the first compressor and the mid-span axial            location on the leading edge of the most upstream aerofoil            of the second compressor; and    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   wherein a fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

is equal to or greater than 0.125.

The fan diameter ratio may be less than or equal to 0.17.

The fan diameter may be greater than 240 cm and less than or equal to380 cm, and optionally may be greater than 300 cm and less than or equalto 380 cm.

The fan diameter may be between 330 cm and 380 cm, and optionally may bebetween 335 cm and 360 cm.

The number of fan blades may be between 16 and 22.

The gas turbine engine may further comprise a gearbox that receives aninput from a core shaft and outputs drive to the fan so as to drive thefan at a lower rotational speed than the core shaft.

A gear ratio of the gearbox may be between 3.1 and 4.0.

The first flange connection may be at, or axially downstream of, aleading edge of the most upstream aerofoil of the second compressor.Alternatively, the first flange connection may be at, or axiallyupstream of, a leading edge of the most upstream aerofoil of the secondcompressor.

The engine may further comprise: a first turbine, and a first core shaftconnecting the first turbine to the first compressor; and a secondturbine and a second core shaft connecting the second turbine to thesecond compressor. The second turbine, second compressor, and secondcore shaft may be arranged to rotate at a higher rotational speed thanthe first core shaft.

The engine core further comprises an inner core casing provided radiallyinwardly of the compressor blades of the compressor system, the innercore casing and the outer core casing defining a core working gas flowpath therebetween. A gas path radius is defined as the outer radius ofthe core gas flow path at the axial position of the first flangeconnection. A gas path ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$

may be equal to or greater than 1.10 and less than or equal to 2.0.

A fan blade mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

may be equal to or less than 19.0 mm/lb.

A blade set mass ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{total}\mspace{14mu}{mass}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{plurality}\mspace{14mu}{of}\mspace{14mu}{fan}\mspace{14mu}{blades}}$

may be in the range between 0.95 mm/lb and 0.35 mm/lb.

The gas turbine engine may further comprise a front mount arranged to beconnected to a pylon.

A front mount position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.18.

A front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.145.

The front mount may be a core mount.

The gas turbine engine may further comprise: a nacelle surrounding theengine core and defining a bypass duct between the engine core and thenacelle; and a fan outlet guide vane (OGV) extending radially across thebypass duct between an outer surface of the engine core and the innersurface of the nacelle, the fan OGV having a radially inner edge and aradially outer edge, wherein an axial midpoint of the radially inneredge is defined as the fan OGV root centrepoint.

A fan OGV root position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 2.6.

A fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.33.

A fan OGV tip position ratio of:

$\frac{\begin{matrix}{{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$

may be equal to or less than 1.8.

A fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{14mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

may be less than or equal to 0.22.

According to a further aspect, there is provided a gas turbine for anaircraft comprising an engine core comprising a compressor system withcompressor blades comprising respective aerofoils, the compressor systemcomprising a first, lower pressure, compressor, and a second, higherpressure, compressor, an outer core casing surrounding the compressorsystem. The gas turbine further comprises a fan located upstream of theengine core, the fan comprising a plurality of fan blades and having afan diameter.

The outer core casing comprises a first flange connection arranged toallow separation of the outer core casing at an axial position of thefirst flange connection, the first flange connection having a firstflange radius, wherein the first flange connection is the first flangeconnection that is downstream of an axial position defined by the axialmidpoint between the mid-span axial location on the trailing edge of themost downstream aerofoil of the first compressor and the mid-span axiallocation on the leading edge of the most upstream aerofoil of the secondcompressor, and a front mount arranged to be connected to a pylon.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

is equal to or greater than 0.125 and less than or equal to 0.17.

Any of the features of the preceding aspect may apply in variousembodiments.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising a first flange connection arranged to allow            separation of the outer core casing at an axial position of            the first flange connection, the first flange connection            having a first flange radius, wherein the first flange            connection is the first flange connection that is downstream            of an axial position defined by the axial midpoint between            the mid-span axial location on the trailing edge of the most            downstream aerofoil of the first compressor and the mid-span            axial location on the leading edge of the most upstream            aerofoil of the second compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially inner edge is defined as the fan OGV root        centrepoint.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

is equal to or greater than 0.125 and less than or equal to 0.17.

Any of the features of the preceding aspect may apply in variousembodiments.

According to a further aspect, there is provided a gas turbine enginefor an aircraft comprising:

-   -   an engine core comprising:        -   a compressor system with compressor blades comprising            respective aerofoils, the compressor system comprising a            first, lower pressure, compressor, and a second, higher            pressure, compressor; and        -   an outer core casing surrounding the compressor system and            comprising a first flange connection arranged to allow            separation of the outer core casing at an axial position of            the first flange connection, the first flange connection            having a first flange radius, wherein the first flange            connection is the first flange connection that is downstream            of an axial position defined by the axial midpoint between            the mid-span axial location on the trailing edge of the most            downstream aerofoil of the first compressor and the mid-span            axial location on the leading edge of the most upstream            aerofoil of the second compressor;    -   a fan located upstream of the engine core, the fan comprising a        plurality of fan blades and having a fan diameter;    -   a nacelle surrounding the engine core and defining a bypass duct        between the engine core and the nacelle; and    -   a fan outlet guide vane (OGV) extending radially across the        bypass duct between an outer surface of the engine core and the        inner surface of the nacelle, the fan OGV having a radially        inner edge and a radially outer edge, wherein an axial midpoint        of the radially outer edge is defined as the fan OGV tip        centrepoint.

A fan diameter ratio of:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{fan}\mspace{14mu}{diameter}}$

-   -   is equal to or greater than 0.125 and less than or equal to        0.17.

Any of the features of the preceding aspect may apply in variousembodiments.

The skilled person would appreciate that larger engines may presentseveral structural challenges; one of which may be managing the bendingstiffness at the engine core to react the increased bending loadsarising from the larger fan diameter. The increased bending load may beexacerbated by fan outlet guide vanes moving rearward as compared toprior engine designs (due to other design constraints or preferences),so making the core bending more significant as the engine gets larger.The changes in relative component positions as detailed in the variousaspects above may help to increase engine core stiffness, and/or improvethe ability of the engine core to react the increased bending loadswithout deforming.

The skilled person would appreciate that these structural challenges maybe particularly relevant to a medium to large gas turbine engine (largerthan 240 cm (95″)—medium—or larger than 300 cm (120″)—large—fandiameter) having a reduction gearbox between its fan and its lowpressure turbine.

As noted elsewhere herein, the present disclosure may relate to a gasturbine engine. Such a gas turbine engine may comprise an engine corecomprising a turbine, a combustor, a compressor, and a core shaftconnecting the turbine to the compressor. Such a gas turbine engine maycomprise a fan (having fan blades) located upstream of the engine core.

Arrangements of the present disclosure may be particularly, although notexclusively, beneficial for fans that are driven via a gearbox.Accordingly, the gas turbine engine may comprise a gearbox that receivesan input from the core shaft and outputs drive to the fan so as to drivethe fan at a lower rotational speed than the core shaft. The input tothe gearbox may be directly from the core shaft, or indirectly from thecore shaft, for example via a spur shaft and/or gear. The core shaft mayrigidly connect the turbine and the compressor, such that the turbineand compressor rotate at the same speed (with the fan rotating at alower speed).

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave any desired number of shafts that connect turbines and compressors,for example one, two or three shafts. Purely by way of example, theturbine connected to the core shaft may be a first turbine, thecompressor connected to the core shaft may be a first compressor, andthe core shaft may be a first core shaft. The engine core furthercomprises a second compressor. The first compressor and a secondcompressor may together be described as forming a compressor system. Thefirst compressor may be a lower pressure compressor that the secondcompressor. The engine core may further comprise a second turbine, and asecond core shaft connecting the second turbine to the secondcompressor. The second turbine, second compressor, and second core shaftmay be arranged to rotate at a higher rotational speed than the firstcore shaft.

The second compressor may be positioned axially downstream of the firstcompressor. The second compressor may be arranged to receive (forexample directly receive, for example via a generally annular duct) flowfrom the first compressor.

In embodiments with a gearbox, the gearbox may be arranged to be drivenby the core shaft that is configured to rotate (for example in use) atthe lowest rotational speed (for example the first core shaft in theexample above). For example, the gearbox may be arranged to be drivenonly by the core shaft that is configured to rotate (for example in use)at the lowest rotational speed (for example only be the first coreshaft, and not the second core shaft, in the example above).Alternatively, the gearbox may be arranged to be driven by any one ormore shafts, for example the first and/or second shafts in the exampleabove.

The gearbox may be a reduction gearbox (in that the output to the fan isa lower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, or 3.2 to 3.8, for example on the order of or at least 3, 3.1,3.2, 3.3, 3.4, 3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratiomay be, for example, between any two of the values in the previoussentence. Purely by way of example, the gearbox may be a “star” gearboxhaving a ratio in the range of from 3.1 or 3.2 to 3.8. In somearrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan and compressors.For example, the combustor may be directly downstream of (for example atthe exit of) the second compressor. By way of further example, the flowat the exit to the combustor may be provided to the inlet of the secondturbine, where a second turbine is provided. The combustor may beprovided upstream of the turbine(s).

Each compressor (for example the first compressor and second compressoras described above—more compressors may be present in other embodiments)may comprise any number of stages, for example multiple stages. Eachstage may comprise a row of rotor blades and a row of stator vanes,which may be variable stator vanes (in that their angle of incidence maybe variable). The row of rotor blades and the row of stator vanes may beaxially offset from each other.

The or each turbine (for example the first turbine and second turbine asdescribed above) may comprise any number of stages, for example multiplestages. Each stage may comprise a row of rotor blades and a row ofstator vanes. The row of rotor blades and the row of stator vanes may beaxially offset from each other.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 0.28 to 0.32. These ratios maycommonly be referred to as the hub-to-tip ratio. The radius at the huband the radius at the tip may both be measured at the leading edge (oraxially forwardmost) part of the blade. The hub-to-tip ratio refers, ofcourse, to the gas-washed portion of the fan blade, i.e. the portionradially outside any platform.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 220 cm, 230 cm, 240 cm, 240 cm, 250 cm (around 100inches), 260 cm, 270 cm (around 105 inches), 280 cm (around 110 inches),290 cm (around 115 inches), 300 cm (around 120 inches), 310 cm, 320 cm(around 125 inches), 330 cm (around 130 inches), 340 cm (around 135inches), 350 cm, 360 cm (around 140 inches), 370 cm (around 145 inches),380 (around 150 inches) cm, 390 cm (around 155 inches), 400 cm, 410 cm(around 160 inches) or 420 cm (around 165 inches). The fan diameter maybe in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 240 cm to 280 cm or 330 cm to 380 cm. Thefan diameter may be greater than 240 cm—an engine comprising a fan withsuch a diameter may be classed as a medium or large engine. The fandiameter may be greater than 300 cm—an engine comprising a fan with sucha diameter may be classed as a large engine. For example, the fandiameter of a large engine may be between 330 cm and 380 cm, andoptionally between 335 cm and 360 cm.

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 220 cm to 300 cm (for example 240 cm to 280 cm, or 250 cmto 270 cm) may be in the range of from 1700 rpm to 2500 rpm, for examplein the range of from 1800 rpm to 2300 rpm, for example in the range offrom 1900 rpm to 2100 rpm. Purely by way of further non-limitativeexample, the rotational speed of the fan at cruise conditions for anengine having a fan diameter in the range of from 330 cm to 380 cm maybe in the range of from 1200 rpm to 2000 rpm, for example in the rangeof from 1300 rpm to 1800 rpm, for example in the range of from 1400 rpmto 1800 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity Utip. The work done by the fan blades13 on the flow results in an enthalpy rise dH of the flow. A fan tiploading may be defined as dH/U_(tip) ², where dH is the enthalpy rise(for example the 1-D average enthalpy rise) across the fan and U_(tip)is the (translational) velocity of the fan tip, for example at theleading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.28, 0.29,0.3, 0.31, 0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (allunits in this paragraph being Jkg⁻¹/(ms⁻¹)²). The fan tip loading may bein an inclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds), for examplein the range of from 0.28 to 0.31 or 0.29 to 0.3.

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds), forexample in the range of from 13 to 16, or 13 to 15, or 13 to 14. Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall pressure ratio of a gas turbine engine as described and/orclaimed herein may be defined as the ratio of the stagnation pressureupstream of the fan to the stagnation pressure at the exit of thehighest pressure compressor (before entry into the combustor). By way ofnon-limitative example, the overall pressure ratio of a gas turbineengine as described and/or claimed herein at cruise may be greater than(or on the order of) any of the following: 35, 40, 45, 50, 55, 60, 65,70, 75. The overall pressure ratio may be in an inclusive range boundedby any two of the values in the previous sentence (i.e. the values mayform upper or lower bounds), for example in the range of from 50 to 70.

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹s, 105 Nkg⁻¹s, 100 Nkg⁻¹s, 95 Nkg⁻¹s, 90 Nkg⁻¹s, 85 Nkg⁻¹s or 80Nkg⁻¹s. The specific thrust may be in an inclusive range bounded by anytwo of the values in the previous sentence (i.e. the values may formupper or lower bounds), for example in the range of from 80 Nkg⁻¹s to100 Nkg⁻¹s, or 85 Nkg⁻¹s to 95 Nkg⁻¹s. Such engines may be particularlyefficient in comparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). Purely by way of example, agas turbine as described and/or claimed herein may be capable ofproducing a maximum thrust in the range of from 330 kN to 420 kN, forexample 350 kN to 400 kN. The thrust referred to above may be themaximum net thrust at standard atmospheric conditions at sea level plus15 degrees C. (ambient pressure 101.3 kPa, temperature 30 degrees C.),with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds), for example in the range of from 1800K to 1950K. The maximumTET may occur, for example, at a high thrust condition, for example at amaximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades. The number of fan blades may be between 16 and 22.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of the gas turbine engine that provides athrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided an aircraft comprising a gasturbine engine as described and/or claimed herein. The aircraftaccording to this aspect is the aircraft for which the gas turbineengine has been designed to be attached. Accordingly, the cruiseconditions according to this aspect correspond to the mid-cruise of theaircraft, as defined elsewhere herein.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine;

FIG. 2 is a close up sectional side view of an upstream portion of a gasturbine engine;

FIG. 3A is a partially cut-away view of a gearbox for a gas turbineengine;

FIG. 3B shows a schematic view of flange connections of a gas turbineengine;

FIG. 4A shows a schematic view of an engine;

FIG. 4B shows a shear force diagram corresponding to FIG. 4A;

FIG. 4C shows a bending moment diagram corresponding to FIG. 4A;

FIG. 5 is a close up sectional side view of an upstream portion of a gasturbine engine with an intercase portion highlighted;

FIG. 6A is a close up sectional side view of an intercase portion;

FIG. 6B is a close up sectional side view of a different intercaseportion;

FIG. 7 is a close up sectional side view of a portion of a gas turbineengine behind the fan, with component spacings and radii marked;

FIG. 8 is an enlarged view of a portion of FIG. 7;

FIG. 9 shows an aircraft with two engines mounted thereon;

FIG. 10 is a schematic sectional side view of the mounting of an engineto a wing of the aircraft;

FIG. 11A is a schematic sectional side view illustrating first flangeposition in one embodiment; and

FIG. 11B is a schematic sectional side view illustrating first flangeposition in another embodiment.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

An exemplary arrangement for a geared fan gas turbine engine 10 is shownin FIG. 2. The low pressure turbine 19 (see FIG. 1) drives the shaft 26,which is coupled to a sun wheel, or sun gear, 28 of the epicyclic geararrangement 30. Radially outwardly of the sun gear 28 and intermeshingtherewith is a plurality of planet gears 32 that are coupled together bya planet carrier 34. The planet carrier 34 constrains the planet gears32 to precess around the sun gear 28 in synchronicity whilst enablingeach planet gear 32 to rotate about its own axis. The planet carrier 34is coupled via linkages 36 to the fan 23 in order to drive its rotationabout the engine axis 9. Radially outwardly of the planet gears 32 andintermeshing therewith is an annulus or ring gear 38 that is coupled,via linkages 40, to a stationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3A. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3A. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIG. 2. For example, where the gearbox 30 has a stararrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example. In somearrangements, the gas turbine engine 10 may not comprise a gearbox 30.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

The engine 10 can be subject to bending due to both static and dynamicloading conditions. A simplified engine bending scenario is show inFIGS. 4A, 4B and 4C. In the embodiment being described, the engine 10 isa medium-large, geared gas turbine engine 10, having a fan diameter 112greater than 240 cm, and more particularly greater than 300 cm. Theengine 10 of the embodiment being described may therefore be describedas a large engine, for example having a fan diameter 112 between 330 cmand 380 cm, and optionally between 335 cm and 360 cm, a gear ratiobetween 3.1 and 4.0, and a number of fan blades between 16 and 22.

A schematic side view of the gas turbine engine 10 is shown in FIG. 4A,with the rotational axis 9 extending horizontally. FIG. 4B shows adiagram of shear force with distance along the rotational axis 9,aligned with the schematic engine view of FIG. 4A. FIG. 4C shows adiagram of bending moment with distance along the rotational axis,aligned with the schematic engine view of FIG. 4A.

Arrow X in FIG. 4A indicates intake upload on the fan 23. The skilledperson would appreciate that take-off is generally the most severecondition where the intake upload is a maximum.

Arrow Y in FIG. 4A indicates reaction load at the front mount 50 due tointake upload. The skilled person would appreciate that gas turbineengines 10 are generally mounted to the wing 52 of an aircraft 90 by oneor more pylons 53, as illustrated in FIGS. 9 and 10. The pylons 53 maybe secured to the engine core 11, to the nacelle 21, or to both. The oreach pylon 53 may be secured to the engine 10 at multiplepoints—multiple mounts may therefore be provided for the pylon(s). Thefront mount 50 is the forward-most mount on the engine 10, and may belocated on the core 11 or on the nacelle 21 in various embodiments. Therear mount is the rear-most mount on the engine 10, and may be locatedon the core 11 or on the nacelle 21 in various embodiments. In FIG. 4A,the front and rear mount are both indicated to be on the engine core11—one or both may instead be on the nacelle 21 in other embodiments.

Arrow Z in FIG. 4A indicates reaction load at the rear mount. In FIG.4A, the front and rear mounts are both indicated to be on the core 11.In alternative embodiments, such as that shown in FIG. 10, the frontmount 53 a may be on the nacelle 21, and the rear mount 53 b may be onthe core 11. In such embodiments, multiple mounts may be present on thecore 11—the mount 53 b may be referred to as the front core mount as itis the forward-most (or only) mount present on the core 11.

The engine core 11 is therefore designed to react the bending momentwith sufficient resistance to reduce or minimise performance losses dueto casing deformations. The skilled person would appreciate thatincreased deformations leads to increased tolerances being needed, suchas an increased blade tip to casing gap, so potentially resulting indecreased efficiency. Additionally or alternatively, casing deformationsmay result in increased wear on bearings, joints and the like, sopotentially reducing engine lifespan.

The skilled person would appreciate that the structural load path of agas turbine engine 10 generally comprises bearing structures, which arerelatively high in stiffness, and rotor/combustor casings, which arerelatively weaker. Flanges that join bearing structures to casings,and/or casing portions to other casing portions, are therefore likely tobe areas where significant changes in stiffness occur. Regionscontaining one or more flanges may therefore be regions where the slope(dw/dl—delta deformation over delta length) tends to be severe.

The casing surrounding the core 11 is arranged to be separated at one ormore positions along its length. A flange connection may be provided toallow separation of the casing into different portions. The positioningof such a flange connection may be constrained by flange integrityconsiderations. In design studies it has been observed that enginestiffness can be improved by moving a flange connection provided toconnect portions of the casing (referred to as the first flangeconnection 60 in the embodiments described herein) further from theengine axis 9—i.e. to a higher diameter relative to the gas path.

Referring to FIG. 5, the low pressure compressor 14 and the highpressure compressor 15 together form a compressor system. The compressorsystem is shown in FIG. 5, and in the close up views of FIGS. 6A and 6B.

Each compressor 14, 15 of the compressor system comprises a respectiveaxial compressor having one or more compressor stages, in theembodiments being described. In alternative embodiments, one or morecentrifugal compressors may be used. In the embodiments being described,each compressor stage comprises a rotor and a stator. In the describedembodiment, each of the high pressure compressor 15 and the lowerpressure compressor 14 comprise two stages formed by a respective firstrotor 62 a, 62 b, first stator 64 a, 64 b, second rotor 66 a, 66 b andsecond stator 68 a, 68 b. Each of the rotors provided in the compressors14, 15 are formed from an annular array of rotor blades arranged torotate in order to provide compression of airflow through the engine 10.Each of the stators comprises an annular array of stator blades that arestationary. The rotor blades and stator blades can each be described asaerofoils provided in the compressors 14, 15.

In the described embodiment two stages are provided in each compressor14, 15. In other embodiments, any other suitable number of stages may beprovided such as a single stage or three or more stages. The number ofstages in each compressor may be the same, as illustrated, or differentfrom each other.

The engine core 11 further comprises a radially inner core casing 70,which is provided radially outwardly of the interconnecting shafts 26,27 connecting the low and high pressure compressors 14, 15 to therespective low and high pressure turbines 17, 19. The inner core casing70 is provided radially inwardly of the blades of the compressors 14,15. The inner core casing 70 extends in a generally axial directionbetween an inlet 72 downstream of the fan 23 and upstream of the lowpressure compressor 15 to an outlet 74 downstream of the high pressurecompressor 15 and upstream of the combustion equipment 16.

The engine core 11 further comprises an outer core casing 76 thatgenerally surrounds the compressor system. The outer core casing 76 isprovided radially outwardly of the inner core casing 70 and the tips ofthe stators and rotors provided in the compressors 14, 15. The coreairflow path A is defined between a radially outer surface of the innercore casing 70 and a radially inner surface of the outer core casing 76.The engine outer core casing 76 extends between the inlet 72 and theoutlet 74 similarly to the inner core casing 70.

The outer core casing 76 comprises a single wall in a forward region ofthe engine 10, and a first outer core casing 78 and a second outer corecasing 80 in a rearward region of the engine 10, in the embodiment beingdescribed. As can be seen in FIG. 5 and the close up view of FIG. 6A,the outer core casing 76 bifurcates into the first and second outer corecasings 78, 80 at a point along its axial length downstream (rearward)of the low pressure compressor 14 and upstream (forward) of the highpressure compressor 15. The first and second outer core casings 78, 80are spaced apart by a gap extending along the axis 9. In the describedembodiment, therefore, only part of the axial length of the outer corecasing 76 is formed from the first and second outer core casings 78, 80.In other embodiments, separate first and second core casings 78, 80could also extend across the low pressure compressor 14, and optionallyacross the full length of the outer core casing 76, or a single wallouter core casing 76 may extend the full length.

The first outer core casing 78 is provided radially inwardly of thesecond outer core casing 80. The inner surface of the first outer corecasing 78 forms the inner surface of the outer core casing 76 whichcontains gas flow within the core airflow A. The first and second outercore casings 78, 80 each provide a separate function within the engine10. The first outer core casing 78 is adapted to contain the coreairflow A. It may therefore be wholly annular and is generally airtight(save for access for bleed ports or the like). The second outer corecasing 80 is instead adapted to provide structural support (i.e. it mayprovide only structural support). It may not therefore need to be whollyannular or airtight. In other embodiments, both pressure containment andstructural support may be provided by both the first and second corecasings 78, 80.

The first outer core casing 78 extends radially inwardly in a downstreamdirection towards the engine centreline 9 in a part of the core 11between the low pressure compressor 14 and the high pressure compressor15 (e.g. in a diffuser section between the compressors 14, 15). Thesecond outer core casing 80 on the other hand is relatively straight,and extends radially inwardly in a downstream direction to a lesserextent than the first core casing 78. As can be seen in the close up ofFIG. 6A, downstream of the point at which the outer core wall 76 splitsinto the first and second outer core casings 78, 80, the first outercore casing extends radially inwardly in a downstream direction to agreater extent (i.e. at a steeper angle towards the engine centreline)than the second outer core casing 80. This results in an annularinter-casing gap 82 being defined by a radially outer surface of thefirst outer core casing 78 and a radially inner surface of the secondouter core casing 80. In the embodiment being described, thisarrangement of the first and second outer core casings 78, 80 maytherefore provide narrowing of the core airflow path A without narrowingthe outer surface of the outer core casing 76 to the same extent.

In an alternative embodiment, as illustrated in FIG. 6B, the outer corecasing 76 does not bifurcate such that first outer core casing 78 andsecond outer core casing 80 are not present. In this embodiment, thecompressors 14, 15 are surrounded by a single casing formed by the outercore casing 76. In such embodiments, the single wall 76 may increase inwidth and/or change shape at a point along its axial length upstream ofthe low pressure compressor 14 and downstream of the high pressurecompressor 15.

First Flange Connection

The first flange connection 60 forms a connection at one end region ofthe “intercase” 76 b of the engine 10—i.e. a part of the outer corecasing 76 between the casing 76 a of the low pressure compressor 14 andthe casing 76 c of the high pressure compressor 15, as illustrated inFIG. 3B.

In the embodiment being described, the first flange connection 60comprises two flanges 60 a, 60 b that extend radially outward fromadjacent portions of the outer core casing 76, and which extendcircumferentially around the casing 76. The two flanges of the firstflange connection 60 extend radially outward from the second outer corecasing 80 in the embodiment shown in FIG. 6A, and radially outward fromthe single wall outer casing 76 in the embodiment shown in FIG. 6B. Inalternative embodiments, the first flange connection 60 may comprise asingle flange arranged to be connected to a connection block, hollowportion of the casing 76 or the likes, instead of to a second flange.The first flange connection 60 may therefore comprise one or moreflanges.

The intercase 76 b may be arranged to be removable or detachable so asto allow access to the first and second compressors 14, 15.

The first flange connection 60 is arranged to allow separation of theouter core casing 76 at the axial position of the first flange 60connection, for example to facilitate access for servicing andmaintenance—the first flange connection 60 therefore defines aseparation point of the engine 10. Two portions 10 a, 10 b of the casing76 of the engine 10 may be separated by disconnection of the firstflange 60 connection (where portion 10 a may correspond to the lowpressure compressor casing 76 a and the intercase 76 b, and portion 10 bto the high pressure compressor casing 76 c, in the examples shown inFIGS. 3B and 6B).

The first flange connection 60 comprises a two-part connection formed bya flange 60 a and a respective connection structure 60 b (i.e. anotherflange, bulkhead, or other structure) to which the flange 60 a isconnected. In the embodiment being described, the flange 60 a of thefirst flange connection 60 is a flange extending from the intercase 76b, and the connection structure 60 b is a flange extending from thecasing 76 c of the high pressure compressor 15. In the embodiment beingdescribed, the flange 60 a of the first flange connection 60 is therearmost flange of the intercase 76 b; in alternative embodiments, a orthe flange forming a part of the first flange connection 60 may beintegral with the intercase 76 but not the rearmost flange of theintercase, may be integral with the casing 76 a of the low pressurecompressor (e.g. being the most downstream flange of the low pressurecompressor casing 76 a), or may be integral with the casing 76 c of thehigh pressure compressor (e.g. being the most upstream flange of thehigh pressure compressor casing 76 c).

The axial position of the first flange connection 60 is defined as theaxial position of the contact surface of the one or more flanges 60 a,60 b from which it is formed. The axial position therefore correspondsto the axial position of the separation point formed by the first flangeconnection 60.

For example, in one embodiment, the first flange connection 60 is formedby a pair of cooperating flanges 60 a, 60 b via which the two portions10 a, 10 b are connected. An example of this is shown in FIGS. 3B and 6Aand described in more detail later. In this embodiment, the axialposition of the first flange connection 60 is defined as the axialposition of the contact surface at which one of the pair of flanges isconnected to, and in contact with, the other.

In other embodiments, the first flange connection 60 comprises a singleflange 60 a that is connected to another structure such as a bulkhead,box-portion or similar structure. An example of this is shown in FIG. 6Band described in more detail later. In this embodiment, the axialposition of the first flange connection 60 is defined as the axialposition of the contact surface of the single flange 60 a from which thefirst flange connection 60 is formed.

In the embodiment being described, the first flange 60 a of the firstflange connection 60 forms part of a first engine casing portion 10 a,and is connected to a second engine casing portion 10 b by a flangeconnector 61.

The two parts 60 a,b of the first flange connection 60 are connected bya flange connector 61. In the embodiment being described the flangeconnector 61 comprises a plurality of bolts passing through the firstflange 60 a of the first flange connection 60 and into a second opposingflange 60 b provided on the second engine casing portion 10 b. In thisembodiment, the first flange 60 a comprises a plurality of holestherethrough arranged to receive the bolts 61, with corresponding holesprovided in the second flange 60 b. In alternative embodiments, one ormore clamps, clips and/or fasteners may be used in addition to, orinstead of, bolts 61. In such embodiments, the first and/or secondflange 60 a,b may not have holes therethrough. In other embodiments, thebolts may pass through holes provided in a single flange 60 a formingthe first flange connection 60 into a bulkhead or other structure towhich the flange is connected.

The first flange connection 60 is the first flange connection that isdownstream of an axial position, X₂, defined by the axial midpointbetween the mid-span axial location, X₁, on the trailing edge of themost downstream low pressure aerofoil of the low pressure compressor 14(the first compressor 14) and the mid-span axial location, X₃, on theleading edge of the most upstream high pressure aerofoil of the highpressure compressor 15 (the second compressor 15). I.e. it is the flange60 connection closest to that axial midpoint, X₂, in a downstreamdirection (as marked by arrow C in FIGS. 11A and 11B) from the axialmidpoint, X₂, the axial midpoint being the midpoint between the rear ofthe forwardmost compressor 14 and the front of the rearmost compressor15 in the embodiments being described.

The skilled person would appreciate that flange connection arrangementsmay vary in various embodiments. For example, in some embodiments thefirst flange connection 60 may be the first flange connection downstreamof the first compressor 14, whereas in other embodiments an additionalone or more flange connections 63 may be present between the firstcompressor 14 and the first flange 60 connection, and/or downstream(rearward) of the first flange 60 connection.

In various embodiments, the additional flange connections 63 may belocated anywhere along the length of a casing of the (first) lowpressure compressor 14. In some embodiments, the additional flangeconnection 63 is located downstream of the first compressor 14. In someengine designs, for example, presence of a core mount 53 b connectingthe core 11 to the pylon and torque box may necessitate a joint in thecore casing at the start of the torque box support structure (rearwardof the first compressor 14). There may be no barrel-shaped casingextending along the length of the first compressor 14 to meet adifferent compressor casing and/or forward support structure.

In other embodiments, the additional flange connection 63 may be locatedat a position along the axial length of first compressor 14. In someengine designs, for example, where the only mount(s) provided may be tothe nacelle 21 rather than to the core 11, no torque box or torque panelmay be provided within the engine core 11—in such embodiments, thecasing may extend further forward—for example to half way along thelength of the first compressor 14.

In some embodiments, the additional flange connection 63 may not bepresent. In such an embodiment, the low pressure compressor casing 76 aand the intercase 76 b may form a single casing rather than being splitinto separate sections. The low pressure compressor casing 76 b thenextends up to, and is connected to, the high pressure compressor casing76 c (e.g. via the first flange connection 60).

In alternative embodiments, such as that shown in FIG. 11A, the firstflange connection 60 is axially upstream of a leading edge of a first(or most upstream) aerofoil of the second compressor 15.

In some embodiments, such as that shown in FIG. 11B, the first flangeconnection 60 is axially downstream of a leading edge of a firstaerofoil of the second compressor 15. In the embodiment of FIG. 11B, thefirst flange connection 60 has an axial position part way along thesecond compressor 15. In alternative embodiments, the first flangeconnection 60 may be axially aligned to the mid-span leading edge of themost upstream high pressure aerofoil of the high pressure compressor 15.

In FIGS. 11A and 11B, the span 103 between X₁ and X₃ is marked. However,this span only serves to define the axial position X₂, and does notlimit the position of the flange connection 60—the flange connection 60may be anywhere downstream of the axial position X₂, as illustrated byarrow C.

In the embodiment being described, the intercase 76 b comprises twoflanges—a forward flange nearer the first compressor 14 and a rearwardflange 60 a nearer the second compressor 15. The two flanges may eachform a part of a different flange connection, and may allow an intercaseportion 76 b of casing 76 to be lifted away to facilitate access to thecompressors 14, 15. In the embodiment being described, the rearwardflange 60 a of the intercase forms part of the first flange connection60 (as the forward flange lies forward of the axial midpoint X₂). Inalternative embodiments, the intercase 76 b may be divided into two ormore portions, and/or a larger number of flanges may be present—thefirst flange 60 a may therefore not be the rearward, or rearmost, flangeof the intercase portion 76 b in all embodiments.

In the embodiment illustrated in FIG. 6A the first flange connection 60is provided in the second outer core casing 80. In this embodiment, thesecond outer core casing 80 is separated into two portions at theseparation point formed by the first flange connection 60. A firstflange 60 a forming the first flange connection 60 is provided on thedownstream of those portions. An opposing second flange 60 b is providedon the other portion of the second outer core casing 80 with which thefirst flange is coupled via the flange connector 61. In otherembodiments, any other suitable structure may be provided to provide aconnecting point for the flange connector 61.

In the embodiment of FIG. 6B, the first flange connection 60 is providedin the outer core casing 76. In this case, the first and second outercore casings are not provided at the axial position of the first flangeconnection 60—the outer core casing 76 instead comprises a single wall.In this embodiment, the first flange connection 60 comprises a singleflange 60 a that is arranged to couple to an adjacent portion of theouter core casing 76, in this embodiment a box-type portion of the outercasing (which may be described as the outer core casing 76 separatinginto first and second outer core casings over a relatively short axiallength of the engine 10). In this embodiment, no second, opposing flangeis provided—the flange connector 61 connects the flange 60 a of thefirst flange connection 60 directly to the opposing casing surface.

In the embodiment being described, the opposing casing surface comprisesthreaded holes arranged to align with threaded holes in the flange 60 a;bolts 61 may then be used to join the flange 60 to the opposing casingsurface.

First Flange Radius

The first flange radius 104 is the radial distance between the enginecentre line 9 and the flange connector 61. In the embodiment beingdescribed, the flange connector 61 comprises a plurality of bolts, andthe first flange radius 104 is defined as the distance between theengine centreline 9 and a centreline of each bolt (the bolts beingoriented axially and located at the same radial distance from the enginecentreline 9).

The skilled person would appreciate that the flange connector location(i.e. bolt location in the embodiment being described) affects stressand strain distribution and may therefore be a more relevant parameterthan the location of the radially outer edge of the first flangeconnection 60.

An increase in first flange radius 104 therefore corresponds to movingthe first flange connection 60 further from the engine centreline 9,and/or moving the flange connector 61 further up the flange provided inthe first flange connection 60 (e.g. by providing bolt holes at a higherradius).

In the embodiments being described, the first flange radius 104 is inthe range of 15 cm to 90 cm, and more particularly in the range from 25cm to 60 cm, for example from 30 cm to 55 cm.

Gas Path Radius and Gas Path Ratio

Referring to FIG. 7, a gas path radius 102 is defined as the outerradius of the core gas flow path A at the axial position of the firstflange connection 60. The gas path radius is measured in the same planeas the first flange radius 104, and is measured from the enginecentreline 9. In the described embodiment, the gas path radius 102 isdefined as the radius of the radially inner surface of the first outercore casing 78 which defines the core gas flow path A measured from theengine centreline 9. In other embodiments, the gas path radius may bemeasured to the radially inner surface of the outer core casing 76 whichdefines the core has flow path A (e.g. in embodiments where the outercore casing 76 is not bifurcated into the first and second outer corecasings 78, 80 at the position of the first flange connection 60).

A gas path ratio is defined as:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}{{gas}\mspace{14mu}{path}\mspace{14mu}{radius}\mspace{14mu}(102)}$

In the embodiment being described, the gas turbine engine 10 isconfigured such that the gas path ratio is equal to or greater than1.10, and more particularly equal to or greater than 1.50. In bothcases, the gas path ratio may be less than 2.0. It may therefore be inan inclusive range between 1.10 and 2.0 or in an inclusive range between1.50 and 2.0.

The radial positioning of the first flange connection 60 relative to theradius of the gas flow path may contribute to reducing or minimisingengine bending whilst maintaining flange integrity. By configuring thegas turbine engine 10 so that the gas path ratio is within the rangeabove the appropriate stiffness may be provided to the engine core 11.

The gas path ratio may be equal to or greater than 1.10 for a mediumsized engine (i.e. fan diameter 112 greater than 240 cm). The gas pathratio may be equal to or greater than 1.50 for a large sized engine(i.e. fan diameter 112 greater than 300 cm). These values may however beassociated with other fan sizes.

In various embodiments, the gas path ratio may have a value of 1.10,1.15. 1.20, 1.25, 1.30, 1.35, 1.40, 1.45, 1.50, 1.55, 1.60, 1.65, 1.70,1.75, 1.80, 1.85, 1.90, 1.95 and 2.00. The gas path ratio may be, forexample, between any two of the values in the previous sentence.

Fan Diameter Ratio

As already described elsewhere herein, the gas turbine engine 10comprises a fan 23 located upstream of the engine core 11. The fan 23comprises a plurality of rotor blades 23 a, also referred to as fanblades 23 a, one of which is shown in FIG. 5. The plurality of rotorblades form a rotor blade set in an annular array around a central hub.

A fan diameter ratio is defined as:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}{{fan}\mspace{14mu}{diameter}\mspace{14mu}(112)}$

In the embodiment being described, the gas turbine engine is configuredsuch that the fan diameter ratio is equal to or greater than 0.125, andmore particularly less than or equal to 0.17. It may therefore be in aninclusive range between 0.125 and 0.17.

The fan diameter is equal to twice the radius 101 of the fan 23. In theembodiment being described, the fan diameter is greater than 240 cm, andmore particularly greater than 300 cm (in both cases it may be no morethan a maximum of 380 cm). In the embodiment being described, the fandiameter is between 330 cm and 380 cm, and more particularly between 335cm and 360 cm.

The radial positioning of the first flange connection 60 relative to thefan 23 contributes to reducing or minimising engine bending whilstmaintaining flange integrity. By configuring the gas turbine engine 10so that the fan diameter ratio is within the range above the appropriatestiffness may be provided to the engine core 11.

In various embodiments, the fan diameter ratio may have a value of0.125, 0.130, 0.135, 0.140, 0.145, 0.150, 0.155, 0.160, 0.165 and 0.170.The fan diameter ratio may be, for example, between any two of thevalues in the previous sentence.

Fan Blade Mass and Blade Set Ratio

A fan blade mass ratio is defined as:

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}{{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$

The fan blade mass ratio relates the mass of each fan blade 23 aprovided on the fan 23 to the first flange radius 104. The skilledperson would appreciate that each fan blade 23 a generally has the samemass, within manufacturing tolerances. If the mass of each fan bladediffers significantly, a fan blade mass ratio for each fan blade may bedetermined separately and configured to fall within the ranges definedherein. In the embodiments being described, the gas turbine engine 10 isconfigured such that the fan blade mass ratio is equal to or less than19.0 mm/pound (41.9 mm/kg). More particularly, the fan blade mass ratiois equal to or greater than 5 mm/pound (11 mm/kg) (or 5.0 mm/pound (11.0mm/kg)). It may therefore be in an inclusive range between 19.0 mm/pound(41.9 mm/kg) and 5.0 mm/pound (11.0 mm/kg). The mass of each fan blademay be in a range between 20 lb (9 kg) and 70 lb (32 kg).

In various embodiments, fan blade mass ratio may have a value of 5.0mm/lb (11.0 mm/kg), 6.0 mm/lb (13.2 mm/kg), 7.0 mm/lb (15.4 mm/kg), 8.0mm/lb (17.6 mm/kg), 9.0 mm/lb (19.8 mm/kg), 10.0 mm/lb (22.1 mm/kg),11.0 mm/lb (24.3 mm/kg), 12.0 mm/lb (26.5 mm/kg), 13.0 mm/lb (28.7mm/kg), 14.0 mm/lb (30.9 mm/kg), 15.0 mm/lb (33.1 mm/kg), 16.0 mm/lb(35.3 mm/kg), 17.0 mm/lb (37.5 mm/kg), 18.0 mm/lb (39.7 mm/kg) and 19.0mm/lb (41.9 mm/kg). The blade set mass ratio may be, for example,between any two of the values in the previous sentence.

The radial positioning of the first flange connection 60 (as determinedby the first flange radius 104) and the fan blade mass may alsocontribute to minimising engine bending whilst maintaining flangeintegrity. By configuring the gas turbine engine 10 so that the fanblade mass ratio is within the range above the appropriate stiffness maybe provided to the engine core 11.

A blade set mass ratio is defined as

$\frac{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}{{total}\mspace{14mu}{mass}\mspace{14mu}{of}\mspace{14mu}{the}\mspace{14mu}{plurality}\mspace{14mu}{of}\mspace{14mu}{fan}\mspace{14mu}{blades}}$

The blade set mass ratio relates the total mass of the plurality of fanblades 23 a forming the fan 23 (i.e. the blade set) and the first flangeradius (104). In the embodiments being described, the blade set ratio isthe inclusive range between 0.95 mm/pound (2.09 mm/kg) and 0.35 mm/pound(0.77 mm/kg).

In various embodiments, the blade set mass ratio may have a value of0.35 mm/lb (0.77 mm/kg), 0.40 mm/lb (0.88 mm/kg), 0.45 mm/lb (0.99mm/kg), 0.50 mm/lb (1.10 mm/kg), 0.55 mm/lb (1.21 mm/kg), 0.60 mm/lb(1.32 mm/kg), 0.65 mm/lb (1.43 mm/kg), 0.70 mm/lb (1.54 mm/kg), 0.75mm/lb (1.65 mm/kg), 0.80 mm/lb (1.76 mm/kg), 0.85 mm/lb (1.87 mm/kg),0.90 mm/lb (1.98 mm/kg) and 0.95 mm/lb (2.09 mm/kg). The blade set massratio may be, for example, between any two of the values in the previoussentence.

As discussed elsewhere herein, each of the fan blades 23 a is at leastpartly formed from a metallic material. The metallic material may betitanium based metal or an aluminium based material such as aluminiumlithium alloy.

In other embodiments, each of the fan blades 23 a may be at least partlyformed from a composite material. The composite material may be, forexample, a metal matrix composite and/or an organic matrix composite,such as carbon fibre.

Fan Outlet Guide Vane

A fan outlet guide vane (OGV) 58 is provided that extends radiallyacross the bypass duct 22, between an outer surface of the engine core11 (e.g. the outer core casing 76) and the inner surface of the nacelle21.

The fan outlet guide vane 58 connects the engine core 11 to the nacelle21. The fan OGV 58 may additionally remove or reduce the swirl from theflow coming from the fan 23.

The fan OGV 58 extends between a radially inner edge 58 a (adjacent theengine core 11) and a radially outer edge 58 b (adjacent the nacelle 21)and has a leading (or upstream) edge and a trailing (or downstream) edgerelative to the direction of gas flow B through the bypass duct 22.

An axial position of the radially inner edge 58 a of the OGV 58 isdefined at the axial mid-point of the radially inner edge 58 a. This maybe referred to as the inner axial centrepoint of the OGV 58, or the rootcentrepoint of the OGV 58.

An axial position of the radially outer edge 58 b of the OGV 58 isdefined at the axial mid-point of the radially inner outer edge 58 b.This may be referred to as the outer axial centrepoint of the OGV 58, orthe tip centrepoint of the OGV 58.

The axial distance 108 between the root centrepoint of the OGV 58 a andthe first flange connection 60 is defined as the distance along the axis9 between the axial position of the root centrepoint 58 a of the OGV 58and the axial position of the axial centre point of the first flangeconnection 60. The axial distance 108 between the root centrepoint ofthe OGV 58 a and the first flange connection 60 is less than or equal to135 cm, and more particularly in the range of 30 cm to 130 cm in theembodiment being described. More particularly, it may be in the range of30 cm to 105 cm, more specifically in the range of 50 cm to 105 cm.

The axial distance 110 between the tip centrepoint 58 b of the OGV 58and the first flange connection 60 is defined as the distance along theaxis 9 between the axial position of the tip centrepoint of the OGV 58 band the axial position of the axial centre point of the first flangeconnection 60. The axial distance 110 between the root centrepoint ofthe OGV 58 a and the first flange connection 60 is less than or equal to90 cm, and more particularly in the range of 20 cm to 90 cm in theembodiment being described. Yet more particularly, it may be in therange of 40 cm to 90 cm.

The axial positioning of the fan outlet guide vanes (fan OGVs) 58 mayhave an effect in reducing or minimising engine bending whilstmaintaining flange integrity.

In particular, the engine 10 may be designed such that the axialdistance 108 between the fan OGV root centrepoint 58 a and the firstflange connection 60 is relatively short. A ratio of the axial distance108 between the fan OGV root centrepoint 58 a and the first flangeconnection 60 centre to the first flange radius 104 of 2.6 or less mayprovide an appropriate stiffness for the engine core 11—this ratio maybe referred to as a fan OGV root position ratio, and may be representedas:

$\frac{\begin{matrix}{{{axial}\mspace{14mu}{distance}\mspace{14mu}(108)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{root}\mspace{14mu}{centrepoint}\mspace{14mu}( {58a} )}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}$

In the embodiment being described, the engine 10 is configured such thatthe fan OGV root position ratio has a value of less than or equal to2.6, and more particularly between 2.6 and 0.8 (inclusive).

In various embodiments, the fan OGV root position ratio may have a valueof 2.6, 2.5, 2.4, 2.2, 2.0, 1.8, 1.6, 1.5, 1.4, 1.2, 1.0, or 0.8. Thefan OGV root position ratio may be, for example, between any two of thevalues in the previous sentence.

In some embodiments, a fan OGV root position to fan diameter ratio of:

$\frac{\begin{matrix}{{{axial}\mspace{14mu}{distance}\mspace{14mu}(108)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{root}\mspace{14mu}{centrepoint}\mspace{14mu}( {58a} )}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

is less than or equal to 0.33. The fan diameter is equal to twice theradius 101 of the fan 23. In the embodiment being described, the fandiameter is greater than 240 cm, and more particularly greater than 300cm (in both cases it may be no more than a maximum of 380 cm). In theembodiment being described, the fan diameter is between 330 cm and 380cm, and more particularly between 335 cm and 360 cm.

In the embodiment being described, the engine 10 is configured such thatthe fan OGV root position to fan diameter ratio is greater than or equalto 0.12.

In various embodiments, the fan OGV root position to fan diameter ratiomay have a value of 0.33, 0.32, 0.30, 0.27, 0.25, 0.22, 0.20, 0.17,0.15, or 0.12. The fan OGV root position to fan diameter ratio may be,for example, between any two of the values in the previous sentence.

In some embodiments, the fan OGV root position to fan diameter ratio maytake a value, or fall in a range, as listed above whilst the fan OGVroot position ratio may not take a value, or fall in a range, as listedabove, or vice versa. In other embodiments, both fan OGV root positionratios may take a value, or fall in a range, as listed above.

Additionally or alternatively, the engine 10 may be designed such thatthe axial distance 110 between the fan OGV tip centrepoint 58 b and thefirst flange connection 60 is relatively short. A ratio of the axialdistance 110 between the fan OGV tip centrepoint 58 b and the firstflange connection 60 centre to the first flange radius 104 of 1.8 orless may provide an appropriate stiffness for the engine core 11—thisratio may be referred to as a fan OGV tip position ratio, and may berepresented as:

$\frac{\begin{matrix}{{{axial}\mspace{14mu}{distance}\mspace{14mu}(110)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{tip}\mspace{14mu}{centrepoint}\mspace{14mu}( {58b} )}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}$

In the embodiment being described, the engine 10 is configured such thatthe fan OGV tip position ratio has a value of less than or equal to 1.8,and more particularly between 1.8 and 0.6 (inclusive).

In various embodiments, the fan OGV tip position ratio may have a valueof 1.8, 1.7, 1.6, 1.5, 1.4, 1.3, 1.2, 1.1, 1.0, 0.9, 0.8, 0.7, or 0.6.The fan OGV tip position ratio may be, for example, between any two ofthe values in the previous sentence.

In some embodiments, a fan OGV tip position to fan diameter ratio of:

$\frac{\begin{matrix}{{{axial}\mspace{14mu}{distance}\mspace{14mu}(110)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{tip}\mspace{14mu}{centrepoint}\mspace{14mu}( {58b} )}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$

is less than or equal to 0.22. The fan diameter is equal to twice theradius 101 of the fan 23. In the embodiment being described, the fandiameter is greater than 240 cm, and more particularly greater than 300cm (in both cases it may be no more than a maximum of 380 cm). In theembodiment being described, the fan diameter is between 330 cm and 380cm, and more particularly between 335 cm and 360 cm.

In the embodiment being described, the engine 10 is configured such thatthe fan OGV tip position to fan diameter ratio is greater than or equalto 0.095.

In various embodiments, the fan OGV tip position to fan diameter ratiomay have a value of 0.22, 0.21, 0.20, 0.19, 0.18, 0.17, 0.16, 0.15,0.14, 0.13, 0.12, 0.11, 0.10 or 0.095. The fan OGV tip position to fandiameter ratio may be, for example, between any two of the values in theprevious sentence.

In some embodiments, the fan OGV tip position to fan diameter ratio maytake a value, or fall in a range, as listed above whilst the fan OGV tipposition ratio may not take a value, or fall in a range, as listedabove, or vice versa. In other embodiments, both fan OGV tip positionratios may take a value, or fall in a range, as listed above.

Front Mount

The engine 10 is arranged to be mounted to a wing 52 of an aircraft 90by means of one or more pylons 53 (a pylon may also be referred to as anairframe strut).

In the embodiments being described with respect to FIG. 10, the engine10 is arranged to be connected to a pylon 53 in a minimum of two places.In the embodiment being described, the two places comprise a nacellemount 53 a connecting the nacelle 21 to the pylon 53 and a core mount 53b connecting the core 11 to the pylon 53. The nacelle mount 53 a isforward of the core mount 53 b in this embodiment. The front mount 50 istherefore the nacelle mount 53 a in the embodiment being described.

In the embodiment shown in FIG. 7, the front mount 50 is a core mount,and two core mounts are provided. The front mount 50 is the front coremount 50.

In some embodiments, the front mount 50 may be a nacelle mount 53 a andmay be located at the axial position of the fan OGV tip centrepoint 58b.

In various embodiments, there may be only one core mount, or there maybe multiple core mounts 53 b—for example, the pylon 53 may be connectedto the core 11 in multiple places, or multiple pylons 53 may each beconnected to the core 11.

In various embodiments, there may be only one nacelle mount 53 a, orthere may be multiple nacelle mounts 53 a—for example, the pylon 53 maybe connected to the nacelle 21 in multiple places, or multiple pylons 53may each be connected to the nacelle 21.

The forward-most mount 50, whether it is a nacelle mount 53 a or a coremount 53 b, is defined as the front mount 50.

The axial distance 106 between the front mount 50 and the first flangeconnection 60 is defined as the distance along the axis 9 between theaxial position of the axial centre point of the front mount 50 and theaxial position of the axial centre point of the first flange connection60.

The skilled person would appreciate that the axial positioning of thefront mount 50 may be important for reducing or minimising enginebending whilst maintaining flange integrity. In particular, the engine10 may be designed such that the axial distance 106 between the frontmount 50 and the first flange connection 60 is relatively short toincrease stiffness (in particular increasing intercase stiffness).Keeping the distance 106 relatively short may also improve ease ofassembly and core inspection. In the embodiments being described thefirst flange connection 60 is located at a point where the bendingmoment on the engine core 11 is quite high. The skilled person wouldappreciate that bending moment is generally higher nearer to the frontmount 50. Increasing the first flange radius 104, so providing a largerdiameter for the first flange connection 60, may facilitate reacting therelatively high bending moment.

A ratio of the axial distance 106 between the front mount 50 and thefirst flange connection 60 centre to the first flange radius of 1.18 orless may provide an appropriate stiffness for the engine core 11—thisratio may be referred to as a front mount position ratio, and may berepresented as:

$\frac{\begin{matrix}{{{axial}\mspace{14mu}{distance}\mspace{14mu}(106)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}\mspace{14mu}(50)}\end{matrix}}{{first}\mspace{14mu}{flange}\mspace{14mu}{radius}\mspace{14mu}(104)}$

In the embodiment being described, the engine 10 is configured such thatthe front mount position ratio has a value of less than or equal to1.18, and more particularly between 1.18 and 0.65.

In various embodiments, the front mount position ratio may have a valueof 1.18, 1.14, 1.10, 1.05, 1.00, 0.95, 0.90, 0.85, 0.80, 0.75, 0.70,0.67, or 0.65. The front mount position ratio may be, for example,between any two of the values in the previous sentence.

The axial distance 106 between the first flange connection 60 and thefront mount 50 is between 30 cm and 75 cm in the embodiment beingdescribed, and more particularly around 30 cm.

In some embodiments, a front mount position to fan diameter ratio of:

$\frac{\begin{matrix}{\mspace{14mu}{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}(106)\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{20mu}(60)\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}\mspace{14mu}(50)}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}\mspace{14mu}(112)}$

is less than or equal to 0.145. The fan diameter 112 is equal to twicethe radius 101 of the fan 23. In the embodiment being described, the fandiameter 112 is greater than 240 cm, and more particularly greater than300 cm (in both cases it may be no more than a maximum of 380 cm). Inthe embodiment being described, the fan diameter 112 is between 330 cmand 380 cm, and more particularly between 335 cm and 360 cm.

In the embodiment being described, the engine 10 is configured such thatthe front mount position to fan diameter ratio is greater than or equalto 0.07.

In various embodiments, the front mount position to fan diameter ratiomay have a value of 0.145, 0.140, 0.135, 0.130, 0.125, 0.120, 0.115,0.110, 0.105, 0.100, 0.095, 0.090, 0.085, 0.080, 0.075, or 0.070. Thefront mount position to fan diameter ratio may be, for example, betweenany two of the values in the previous sentence.

In some embodiments, the front mount position to fan diameter ratio maytake a value, or fall in a range, as listed above whilst the front mountposition ratio may not take a value, or fall in a range, as listedabove, or vice versa. In other embodiments, both front mount positionratios may take a value, or fall in a range, as listed above.

In the present disclosure, upstream and downstream are with respect tothe air flow through the compressor system; and front and rear is withrespect to the gas turbine engine, i.e. the fan being in the front andthe turbine being in the rear of the engine.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

The invention claimed is:
 1. A gas turbine engine for an aircraftcomprising: an engine core comprising: a compressor system withcompressor blades comprising respective aerofoils, the compressor systemcomprising a first, lower pressure, compressor, and a second, higherpressure, compressor; and an outer core casing surrounding thecompressor system and comprising a first flange connection arranged toallow separation of the outer core casing at an axial position of thefirst flange connection, the first flange connection having a firstflange radius, the first flange connection being the first flangeconnection that is downstream of an axial position defined by an axialmidpoint between a mid-span axial location on a trailing edge of a mostdownstream aerofoil of the first compressor and a mid-span axiallocation on a leading edge of a most upstream aerofoil of the secondcompressor; a fan located upstream of the engine core, the fancomprising a plurality of fan blades and having a fan diameter; anacelle surrounding the engine core and defining a bypass duct betweenthe engine core and the nacelle; and a fan outlet guide vane (OGV)extending radially across the bypass duct between an outer surface ofthe engine core and an inner surface of the nacelle, the fan OGV havinga radially inner edge and a radially outer edge, an axial midpoint ofthe radially outer edge being defined as a fan OGV tip centrepoint,wherein a fan OGV tip position ratio of: $\frac{\begin{matrix}{{{an}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}{\;\mspace{11mu}}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$is equal to or less than 1.8.
 2. The gas turbine engine of claim 1,wherein the fan OGV tip position ratio is greater than or equal to 0.6.3. The gas turbine engine of claim 1, wherein the fan OGV tip positionratio is less than or equal to 1.20.
 4. The gas turbine engine of claim1, wherein the fan diameter is greater than 240 cm and less than orequal to 380 cm.
 5. The gas turbine engine of claim 1, wherein the fandiameter is between 330 cm and 380 cm.
 6. The gas turbine engine ofclaim 1, wherein a number of the fan blades is between 16 and
 22. 7. Thegas turbine engine of claim 1, further comprising a gearbox that isconfigured to receive an input from a core shaft, and output drive tothe fan so as to drive the fan at a lower rotational speed than the coreshaft, and optionally wherein a gear ratio of the gearbox is between 3.1and 4.0.
 8. The gas turbine engine of claim 1, wherein the first flangeconnection is at, or axially downstream of, the leading edge of the mostupstream aerofoil of the second compressor.
 9. The gas turbine engine ofclaim 1, wherein the first flange connection is at, or axially upstreamof, the leading edge of the most upstream aerofoil of the secondcompressor.
 10. The gas turbine engine according to claim 1, furthercomprising: a first turbine, and a first core shaft connecting the firstturbine to the first compressor; and a second turbine and a second coreshaft connecting the second turbine to the second compressor, whereinthe second turbine, the second compressor, and the second core shaft arearranged to rotate at a higher rotational speed than the first coreshaft.
 11. The gas turbine engine according to claim 1, wherein a fanOGV tip position to fan diameter ratio of: $\frac{\begin{matrix}{{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{11mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{tip}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$ is lessthan or equal to 0.22.
 12. The gas turbine engine according to claim 11,wherein the fan OGV tip position to fan diameter ratio is greater thanor equal to 0.095.
 13. The gas turbine engine of claim 12, wherein thefan OGV root position ratio is greater than or equal to 0.8.
 14. The gasturbine engine according to claim 12, wherein a fan OGV root position tofan diameter ratio of: $\frac{\begin{matrix}{{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{11mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{11mu}{and}\mspace{14mu}{the}\mspace{14mu}{fan}\mspace{14mu}{OGV}\mspace{11mu}{root}\mspace{14mu}{centrepoint}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$ is lessthan or equal to 0.33.
 15. The gas turbine engine according to claim 1,further comprising a front mount arranged to be connected to a pylon,wherein a front mount position ratio of: $\frac{\begin{matrix}{{{an}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}\mspace{14mu}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{20mu}{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}$is equal to or less than 1.18.
 16. The gas turbine engine according toclaim 15, wherein a front mount position to fan diameter ratio of:$\frac{\begin{matrix}{{{the}\mspace{14mu}{axial}\mspace{14mu}{distance}\mspace{14mu}{between}\mspace{14mu}{the}\mspace{14mu}{first}\mspace{14mu}{flange}}\mspace{31mu}} \\{{connection}{\;\mspace{11mu}}{and}\mspace{14mu}{the}\mspace{14mu}{front}\mspace{14mu}{mount}}\end{matrix}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$ is lessthan or equal to 0.145.
 17. The gas turbine engine according to claim 1,wherein: the engine core further comprises an inner core casing providedradially inwardly of the compressor blades of the compressor system, theinner core casing and the outer core casing defining a core working gasflow path therebetween, a gas path radius is defined as an outer radiusof the core working gas flow path at the axial position of the firstflange connection, and a gas path ratio of:$\frac{{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{the}\mspace{14mu}{gas}\mspace{14mu}{path}\mspace{14mu}{radius}}$is equal to or greater than 1.10 and less than or equal to 2.0.
 18. Thegas turbine engine according to claim 1, wherein a fan diameter ratioof:$\frac{{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{{the}\mspace{14mu}{fan}\mspace{14mu}{diameter}}$is equal to or greater than 0.125 and less than or equal to 0.17. 19.The gas turbine engine according to claim 1, wherein a fan blade massratio of:$\frac{{the}\mspace{14mu}{first}\mspace{14mu}{flange}\mspace{14mu}{radius}}{a{\mspace{11mu}\;}{mass}\mspace{14mu}{of}\mspace{14mu}{each}\mspace{14mu}{fan}\mspace{14mu}{blade}}$is equal to or less than 19.0 mm/lb.